A method of assembling a gas turbine engine includes providing at least one heat exchanger assembly including a heat exchanger and a lobed mixer extending downstream from the heat exchanger, wherein the mixer includes a plurality of lobes that each define a first chute and at least one second chute that extends between each pair of adjacent spaced-apart lobes, and coupling the at least one heat exchanger assembly within a bypass duct of the engine such that the at least one heat exchanger assembly is coupled to at least one of an outer engine casing and an inner engine casing of the turbine engine.
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1. A method of assembling a gas turbine engine, said method comprising:
providing at least one heat exchanger assembly including a heat exchanger and a lobed mixer coupled to the heat exchanger and extending downstream from the heat exchanger, wherein the mixer includes a plurality of lobes on multiple sides of the mixer, wherein each lobe define a first chute and at least one second chute that extends between each pair of adjacent spaced-apart lobes; and
coupling the at least one heat exchanger assembly within a bypass duct of the engine such that a first portion of bypass air flows above and below said at least one heat exchanger and the mixer within the bypass duct, while a second portion of bypass air flows through said at least one heat exchanger, wherein the first and second portions of bypass air communicate at said plurality of lobes on said multiple sides of the mixer; wherein the at least one heat exchanger assembly is coupled to at least one of an outer engine casing and an inner engine casing of the turbine engine.
8. A heat exchanger assembly for use in a turbine engine, said heat exchanger assembly comprising:
at least one core heat exchanger configured to be coupled within a bypass duct of the engine, said at least one core heat exchanger including an inlet, an outlet, a hollow body and a plurality of heat exchanger fins that extend at least partially through the hollow body between the inlet and the outlet, said at least one core heat exchanger configured to be positioned within the duct such that a first portion of bypass air flows above and below said at least one core heat exchanger while a second portion of bypass air flows through said at least one core heat exchanger; and
a mixer coupled to said at least one core heat exchanger and extending downstream from said at least one core heat exchanger, said mixer comprising a plurality of lobes on multiple sides of the mixer, wherein each lobe defines a first chute, wherein said plurality of lobes are spaced about an outer perimeter on multiple sides of said mixer such that at least one second chute is defined between each pair of adjacent spaced-apart lobes.
13. A gas turbine engine comprising:
an outer casing,
an inner casing,
a bypass duct defined between said outer casing and said inner casing, the bypass duct defining an airflow path; and
at least one heat exchanger assembly coupled within said bypass duct to at least one of said outer casing and said inner casing, said at least one heat exchanger assembly comprising:
at least one heat exchanger disposed at least partially within said airflow path; and
a mixer coupled to said at least one heat exchanger such that a first portion of bypass air flows above and below said at least one heat exchanger and said mixer within said bypass duct, while a second portion of bypass air flows through said at least one heat exchanger, said mixer comprising a plurality of lobes on multiple sides of said mixer, wherein each lobe defines a first chute, said plurality of lobes spaced about a perimeter of said mixer such that a second chute is defined between each pair of adjacent spaced-apart lobes, wherein the first and second portions of bypass air communicate at said plurality of lobes on said multiple sides of said mixer.
2. A method in accordance with
3. A method in accordance with
4. A method in accordance with
5. A method in accordance with
6. A method in accordance with
7. A method in accordance with
9. A heat exchanger assembly in accordance with
10. A heat exchanger assembly in accordance with
an inlet having a first cross-sectional area;
an outlet having a second cross-sectional area that is larger than said first cross-sectional area; and
a mixer body extending between said inlet and said outlet.
11. A heat exchanger assembly in accordance with
12. A heat exchanger assembly in accordance with
14. A gas turbine engine in accordance with
an inlet, an outlet, and a hollow body extending therebetween; and
a plurality of heat exchanger fins extending at least partially through said body between said inlet and said outlet.
15. A gas turbine engine in accordance with
16. A gas turbine engine in accordance with
an inlet having a first cross-sectional area;
an outlet having a second cross-sectional area that is larger than said inlet cross-sectional area; and
a hollow body extending between said inlet and said outlet.
17. A gas turbine engine in accordance with
18. A gas turbine engine in accordance with
19. A gas turbine engine in accordance with
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The U.S. Government has certain rights in this invention as provided for by the terms of Contract No. N00019-96-C-0176.
This invention relates generally to turbine engines, and more particularly, to heat exchangers used with turbine engines.
At least some known turbine engines include an inlet, a fan assembly, low and high pressure compressors, a combustor, and at least one turbine. Air entering such engines through the inlet is compressed by the fan assembly. In such engines, a portion of the compressed air is channeled through a bypass duct defined generally between an outer engine casing and an inner engine casing. At least some known turbine engines position a heat exchanger within the bypass duct to enable a flow of bypass air to be used by a thermal management system, such as for a heat sink. Specifically, in known engines, a thermal management system provides cooling to various systems that may generate heat by transferring heat from a cooling fluid channeled through the heat exchanger to bypass airflow channeled past the heat exchanger.
At least some known heat exchangers use an inlet scoop to facilitate entraining a sufficient amount of bypass air into the heat exchanger for cooling the various heat-generating systems. However, such heat exchangers may cause undesirable pressure losses in the bypass airflow. In addition, generally such heat exchangers also generate a flow of hot air, or a temperature streak, downstream from the heat exchanger. The flow of hot air and the reduced bypass airflow pressure may adversely reduce the performance and/or efficiency of the turbine engine.
In one aspect a method of assembling a gas turbine engine is provided. The method includes providing at least one heat exchanger assembly including a heat exchanger and a lobed mixer extending downstream from the heat exchanger, wherein the mixer includes a plurality of lobes that each define a first chute and at least one second chute that extends between each pair of adjacent spaced-apart lobes, and coupling the at least one heat exchanger assembly within a bypass duct of the engine such that the at least one heat exchanger assembly is coupled to at least one of an outer engine casing and an inner engine casing of the turbine engine.
In another aspect a heat exchanger assembly for use in a turbine engine is provided. The heat exchanger assembly includes at least one core heat exchanger, and a mixer extending downstream from the at least one core heat exchanger, the mixer comprising a plurality of lobes that each define a first chute, wherein the plurality of lobes are spaced about an outer perimeter of the mixer such that at least one second chute is defined between each pair of adjacent spaced-apart lobes.
In a further aspect a gas turbine engine is provided. The gas turbine engine includes an outer casing, an inner casing, and a bypass duct defined therebetween, at least one heat exchanger assembly coupled within the bypass duct to at least one of the outer casing and the inner casing, and a mixer coupled to the at least one heat exchanger, the mixer comprising a plurality of lobes that each define a first chute, the plurality of lobes spaced about a perimeter of the mixer such that a second chute is defined between each pair of adjacent spaced-apart lobes.
The present invention generally provides exemplary apparatus and methods to facilitate mixing bypass airflow in a gas turbine engine. The embodiments described herein are not limiting, but rather are exemplary only. It should be understood that the present invention may apply to any type of heat exchanger wherein an entrained fluid channeled through a heat exchanger is used to facilitate heat transfer with a fluid channeled past the heat exchanger.
In the exemplary embodiment, engine assembly 10 also includes an outer casing 32 which is radially-spaced from an inner casing 34 including a forward section 36. Outer casing 32 and inner casing 34 define a bypass duct 38. Moreover, at least one heat exchanger assembly 100 is coupled within bypass duct 38. More specifically, in the exemplary embodiment, engine assembly 10 includes four heat exchanger assemblies 100 that are each circumferentially-spaced within bypass duct 38. Alternatively, engine assembly 10 may include any number of heat exchanger assemblies 100 that enables engine assembly 10 to function as described herein.
During operation, air entering engine assembly 10 through intake side 24 is compressed by fan assembly 22. The flow of compressed air exiting fan assembly 22 is split such that a first airflow portion 42 is channeled into core turbine engine 12 and a second airflow portion, or bypass air 44, is channeled through bypass duct 38. First airflow portion 42 is further compressed by high-pressure compressor 14 and is channeled to combustor 16. Airflow discharged from combustor 16 drives rotating turbines 18 and 20 prior to being discharged from engine 10 through exhaust side 26.
Bypass air 44 is channeled generally axially downstream through bypass duct 38 and past heat exchanger assembly 100. Heat exchanger assembly 100 causes bypass air 44 to be split into a first flow of bypass air and a second flow of bypass air. Specifically, in the exemplary embodiment, the first flow of bypass air, or cold bypass air 46, is channeled past heat exchanger assembly 100, and the second flow of bypass air is channeled through heat exchanger assembly 100 such that a flow of hot bypass air 48 is discharged therefrom.
In the exemplary embodiment, each heat exchanger assembly 100 is coupled in communication to a thermal management system (not shown). The thermal management system utilizes bypass air 44 as a heat sink to facilitate cooling heat-generating systems and components of an associated aircraft (not shown) and/or engine 10. Such heat-generating systems may include, but are not limited to only including, aircraft electrical systems, engine cooling systems, and/or cockpit air-conditioning systems. In the exemplary embodiment, as described in more detail below, heat exchanger assembly 100 facilitates transferring heat from a cooling fluid channeled therethrough to the flow of bypass air 44. As a result, the temperature of the cooling fluid is reduced and hot bypass air 48 is discharged from heat exchanger assembly 100.
In the exemplary embodiment, each core heat exchanger 102 is substantially rectangular and includes an inner cavity 120 defined therein. Alternatively, each core heat exchanger 102 may have any shape that is suitable for use within bypass duct 38. In the exemplary embodiment, each core heat exchanger 102 includes a radially outer panel 112, a radially inner panel 114, and a pair of side panels 116 extending between panels 112 and 114. Panels 112, 114, and 116 are coupled together to define cavity 120 therein. Each core heat exchanger 102 also includes cooling tubes (not shown) coupled within cavity 120 to channel cooling fluid therethrough. A plurality of heat transfer fins 122 extend substantially radially inward from inner panel 114 towards outer panel 112, wherein fins 122 are coupled to the cooling tubes. Moreover fins 122, as described in more detail below, increase a surface area of the cooling tubes, which facilitates heat transfer. Each core heat exchanger 102 also includes an inlet side 118 and an outlet side 124. Each heat exchanger assembly 100 is oriented such that inlet side 118 is substantially perpendicular to the flow of bypass air 44. In the exemplary embodiment, as described in more detail below, mixer 104 extends downstream from the outlet side 124 of each core heat exchanger 102, such that mixer 104 is in flow communication with cavity 120.
In the exemplary embodiment, each pair of adjacent core heat exchangers 102 are coupled together via an inlet plenum 126. Specifically, in the exemplary embodiment, inlet plenum 126 is trapezoidal-shaped and facilitates channeling the cooling fluid through the cooling tubes coupled within heat exchanger assembly 100. Alternatively, inlet plenum 126 may have any shape that facilitates coupling the pair of core heat exchangers 102 together in any manner that enables heat exchanger assembly 100 to function as described herein. In the exemplary embodiment, each inlet plenum 126 extends circumferentially between each pair of circumferentially-spaced core heat exchangers 102 within bypass duct 38.
In the exemplary embodiment, each mixer 104 includes an inlet 140, an outlet 142, and a body 144 extending therebetween. Inlet 140 has a cross-sectional shape that is substantially the same as the cross-sectional shape of core outlet side 124. Moreover, inlet 140, in the exemplary embodiment, is coupled against outlet side 124 such that mixer 104 is in flow communication with cavity 120 to enable hot bypass air 48 discharged from core heat exchanger 102 to be channeled towards mixer outlet 142. In the exemplary embodiment, mixer outlet 142 has a cross-sectional area that is larger than the cross-sectional area of inlet 140, as described below in more detail.
Each mixer 104, in the exemplary embodiment, includes a plurality of lobes 146. Specifically, in the exemplary embodiment, each lobe 146 defines a first chute, or hot chute 148 that, as described in more detail below, channels hot bypass air 48 therethrough. In the exemplary embodiment, each pair of adjacent lobes 146 is coupled together by a second or cold chute 150. In an alternative embodiment, more than one cold chute 150 is defined between each pair of hot chutes 148. Each cold chute 150 enables cold bypass air 46 to be channeled therethrough, as described in more detail below.
In the exemplary embodiment, at least one hot chute 148 extends generally divergently from inlet 140 radially outward into the flow of cold bypass air 46, and more specifically, towards at least one of inner casing 34 and/or outer casing 32. Moreover, in the exemplary embodiment, at least one cold chute 150 is generally convergent and extends generally radially inward into the flow of hot bypass air 48, and more specifically, away from at least one of inner casing 34 and outer casing 32. As a result, a lobed or wavy mixer 104 is defined at mixer outlet 142, which as described in more detail below, facilitates entraining bypass air 44 in each core heat exchanger 102, mixing hot bypass air 48 with cold bypass air 46, and reducing the overall pressure losses of bypass air 44. In the exemplary embodiment, each mixer 104 includes eight hot chutes 148 and eight cold chutes 150. Alternatively, mixer 104 may include any number of hot chutes 148 and/or cold chutes 150.
During operation, in the exemplary embodiment, bypass air 44 is channeled generally axially through bypass duct 38 towards each heat exchanger assembly 100. Bypass air 44 is split such that a first portion is channeled through each heat exchanger assembly 100, and a second portion is channeled past each heat exchanger assembly 100. Moreover, the thermal management system transfers heat from the heat-generating systems and components of the aircraft and/or engine 10 to the cooling fluid. Specifically, the thermal management system, in the exemplary embodiment, channels the heated cooling fluid through the cooling tubes coupled within each heat exchanger assembly 100. Fins 122 facilitate increasing the surface area of the cooling tubes, which facilitates reducing a temperature of the cooling fluid channeled therein.
In the exemplary embodiment, the first portion of bypass air 44 is channeled through each heat exchanger assembly 100 to facilitate transferring heat from fins 122, and more specifically, the cooling fluid, to the first portion of bypass air 44 such that hot bypass 48 is discharged from each heat exchanger assembly 100. The second portion of bypass air 44 is channeled past each core heat exchanger 102, and more specifically, through each cold chute 150, such that cold bypass air 46 is discharged from each cold chute 150.
Hot chute 148 channels hot bypass air 48 downstream generally divergently from inlet 140 radially outward into the flow of cold bypass air 46. Moreover, cold chute 150, in the exemplary embodiment, channels cold bypass air 46 downstream generally convergent and generally radially into the flow of hot bypass air 48. The flow of cold bypass air 46 discharged from each cold chute 150 mixes with the flow of hot bypass air 48 discharged from each hot chute 148 downstream from mixer 104. As a result, temperature streaks extending downstream from heat exchanger assembly 100 are facilitated to be reduced.
In the exemplary embodiment, mixer 104 facilitates increasing the amount of bypass air 44 entrained in each heat exchanger assembly 100. Specifically, the cross-sectional area of outlet 142 of mixer 104 facilitates reducing the size of gaps 108 and 110 such that the flow of cold bypass air 46 over mixer 104 is reduced. As a result, the amount of hot bypass air 48 entrained within each core heat exchanger 102 is facilitated to be increased, which facilitates the heat transfer from the cooling fluid channeled through each heat exchanger assembly 100 to hot bypass air 48.
Mixer 104 also facilitates reducing the pressure losses of bypass air 44. In some known heat exchangers, the location of the heat exchanger causes a change in the momentum of the bypass air that results in a loss in pressure of the bypass air. Moreover, in known turbine engines, changing the direction of the flow of the bypass air to a generally circumferential and/or generally radial flow direction causes pressure losses in the bypass air. In the exemplary embodiment, mixer 104 facilitates reducing momentum changes of cold bypass air 46, which facilitates reducing the overall pressure losses of bypass air 44 channeled through bypass duct 38. In the exemplary embodiment, cold bypass air 46 is channeled past heat exchanger 100 and is channeled through cold chute 150. As a result, the flow of cold bypass air 46 is maintained in a generally axial flow direction such that pressure losses of the flow of cold bypass air 46 are facilitated to be reduced.
The above-described methods and apparatus facilitate entraining bypass air in at least one core heat exchanger, mixing a flow of hot bypass air with a flow of cold bypass air, and reducing the overall pressure losses of the bypass air. Specifically, each heat exchanger assembly includes a mixer that includes a plurality of lobes. Each lobe defines a first chute that is configured to channel a hot bypass airflow therethrough. A second chute is defined between each pair of adjacent spaced-apart lobes. Each second chute is configured to channel a cold bypass airflow therethrough. As a result, the flow of hot bypass air discharged from each first chute is mixed with the flow of cold bypass air discharged from each second chute. The cross-sectional area of the mixer outlet facilitates reducing gaps defined between the outer and inner casings and each heat exchanger assembly. As a result of the reduced gaps, the amount of cold bypass air channeled past each heat exchanger assembly is reduced and the amount of bypass air entrained within each heat exchanger assembly is increased. Moreover, each second chute facilitates channeling the cold bypass air in a generally axial direction such that the momentum of cold bypass air is facilitated to be maintained.
Exemplary embodiments of heat exchangers are described in detail above. The mixer described herein is not limited to use with the heat exchangers for the gas turbine engine, but rather, the mixer can be utilized independently and separately from other heat exchanger components described herein. Moreover, the invention is not limited to the embodiments of the mixers described above in detail. Rather, other variations of the mixers may be utilized within the spirit and scope of the claims.
While the invention has been described in terms of various specific embodiments, those skilled in the art will recognize that the invention can be practiced with modification within the spirit and scope of the claims.
Perveiler, David Andrew, MacLean, Malcolm Kenneth, Lcon, Ross Michael
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| Jul 27 2007 | PERVEILER, DAVID ANDREW | General Electric Company | ASSIGNMENT OF ASSIGNORS INTEREST SEE DOCUMENT FOR DETAILS | 019621 | /0246 | |
| Jul 27 2007 | MACLEAN, MALCOLM KENNETH | General Electric Company | ASSIGNMENT OF ASSIGNORS INTEREST SEE DOCUMENT FOR DETAILS | 019621 | /0246 | |
| Jul 27 2007 | LEON, ROSS MICHAEL | General Electric Company | ASSIGNMENT OF ASSIGNORS INTEREST SEE DOCUMENT FOR DETAILS | 019621 | /0246 | |
| Jul 30 2007 | General Electric Company | (assignment on the face of the patent) | / | |||
| Jun 30 2015 | GE AVIATION LEGAL | DEPARTMENT OF THE NAVY | CONFIRMATORY LICENSE SEE DOCUMENT FOR DETAILS | 038837 | /0961 |
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